System and method for implementing a constellation of non-geostationary satellites that does not interfere with the geostationary satellite ring

ABSTRACT

Provided is an improved system and method for implementing a constellation of satellites in inclined elliptical orbits. The satellites are operated during the portion of their orbits near apogee to emulate the characteristics of geostationary satellites. The orbits are configured to form a number of closely spaced repeating ground tracks around the earth. In each ground track the satellites operate only in arcs well above or below the equator to provide a large number of non-geostationary orbital slots that substantially increase global satellite capacity without interfering with the existing geostationary satellite ring. Minimum spacing is maintained between satellites in each active arc and between satellites in the active arcs of adjacent ground tracks to ensure that the satellites in the non-geostationary constellation do not interfere with each other.

REFERENCE TO PRIORITY DOCUMENTS

This Patent Application is a continuation of, and claims priority under35 USC §120 to, co-pending U.S. patent application Ser. No. 11/095,745,entitled “Fixed satellite constellation system employingnon-geostationary satellites in sub-geosynchronous elliptical orbitswith common ground tracks,” filed Mar. 30, 2005, which is a continuationclaiming priority under 35 USC §120 to U.S. application Ser. No.09/658,215, entitled “Fixed satellite constellation system employingnon-geostationary satellites in sub-geosynchronous elliptical orbitswith common ground tracks,” filed Sep. 8, 2000, now U.S. Pat. No.6,954,613, issued Oct. 11, 2005, which claims priority under 35 USC§119(e) to Provisional Patent Application Ser. No. 60/153,289, filed onSep. 10, 1999, all of the above-reference patent applications areincorporated by reference for all purposes.

BACKGROUND

Satellite communications systems often require that a station on theground communicate with a satellite. The satellite tracking issimplified when the satellite appears to be maintained stationaryrelative to the Earth. Geosynchronous (“geo”) satellites have thischaracteristic. However, geo-satellites require high altitude orbits.These high altitude orbits require large payloads and launches, and alsocan have relatively long propagation delays during communication.

SUMMARY

The present disclosure describes an array of non-geostationarysatellites in sub-geosynchronous, inclined elliptical orbits. Each ofthe satellites communicates with a point on the earth. At least aplurality of the satellites is in an elliptical orbit with the earth atone focus of the ellipse.

At and near their apogee points, the satellites move slowly relative tothe Earth. These satellites appear virtually geostationary to userswithin at least part of the desired coverage area.

The disclosed embodiments use three sub-constellations, each with 5satellites. Three total sub-constellations are used. Two of thesesub-constellations are used for Northern Hemisphere operation. A thirdconstellation is for Southern Hemisphere operation. The satellites areactive over only part of their total time of their orbits. The activetime of the orbit is when the satellites are closest to their apogees.

These active times can occur when the satellites are at latitudes above45.degree. These satellites are hence seen at high elevations over muchof their primary service areas.

This system is also effectively transparent to the geostationary fixedsatellite services and can be separated from the geostationary arcpreferably by at least 40.degree. at all times within the service areaof the system.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 shows a basic layout of the multiple elliptical orbits of thepresent invention;

FIG. 1A shows a graphical depiction of the satellite's angular motionalong its orbit as a function of the semi-major axis of the ellipticalorbit.

FIGS. 2A & 2B shows a block diagram of the satellite communicationequipment used according to the present invention;

FIG. 2C shows a flowchart of operation of the satellites of the presentinvention;

FIG. 3 shows the characteristics of a basic ellipse;

FIGS. 4A-4F show characteristics of the three-satellite orbit of thepresent invention;

FIG. 4G shows characteristics of this orbit which prevent interferencewith geosynchronous satellites in an inclined orbit;

FIG. 4H shows characteristics of this orbit which prevent interferencewith geosynchronous satellites in an equatorial orbit;

FIGS. 5A-5E show characteristics of the five satellite orbit of thepresent invention;

FIG. 6 shows an overall view of the ten satellite orbit of the presentinvention;

FIGS. 7A-7G show the positions of the satellites of the ten satelliteembodiment within their repeating ground tracks;

FIG. 8 shows the operating elevation angles for the ten satellite orbit,and their angular isolation from geo satellites; and

FIG. 9 shows ground tracks of the preferred orbits.

DESCRIPTION OF THE PREFERRED EMBODIMENTS

The disclosed system defines a communication system including groundcommunication equipment and a special constellation of satellites inelliptical orbits at lower altitudes than those necessary forgeosynchronous, which simulate to the characteristics of ageosynchronous orbit from the viewpoint of the ground communicationequipment on the earth. The inventors recognized that satellites whichorbit in certain elliptical orbits spend most of their time near theapogees of their orbits: the time when they are most distant from theearth. These satellites spend only a minority of their time near theirperigee. For example, an elliptical satellite in a 12-hour orbit spendseight of those hours near its apogee. By appropriately choosingcharacteristics of the satellite orbit, the satellite can be made toorbit, during that time, at a velocity that approximates the rotationalvelocity of the earth. The disclosed system defines a communicationsystem using a constellation of satellites chosen and operating suchthat a desired point on the earth always tracks and communicates with asatellite at or near apogee.

Another important feature of the disclosed system is the recognition ofhow this mode of operation of the satellite changes its powercharacteristics.

Geosynchronous satellites are used virtually 100% of the time (exceptwhen in eclipse) and hence their power supplies must be capable offull-time powering. This means, for example, if the satellite requires 5Kw to operate, then the power supply and solar cells must be capable ofsupplying a continuous 5 Kw of power. The satellites of the disclosedsystem, however, are not used 100% of the time. During the perigeeportions of the satellite orbit, the satellites are typically not usingmost of their transmit and receive capability and hence, the inventorsrecognized, do not use a large part of their power capability.

The inventors of the disclosed system recognized this feature of thesatellites, and realized that the satellites could be storing the powerthat is being produced during this time of non-use. Therefore, theinventors realized that the size of the power supply could be reduced bya factor of the percentage of time that the satellite is not used.

The power sources can be any known means, including solar cells, nuclearreactors, or the like. If the satellite is used half the time, then thepower source need only be sized to provide half the power. At times whenthe satellite is not being used, the power source provides power to abattery storage cell, which holds the power in reserve for times whenthe satellite is being used.

Like geo systems, the satellite of the disclosed system is virtuallycontinuously in the same location. Unlike geo-based systems, however,the ground communication equipment of the disclosed system does notalways communicate with the same satellite. The satellites move slightlyrelative to the earth, i.e. they are not always precisely at the samepoint in their apogees. One important advantage of the disclosed systemis that the one satellite at apogee later moves to perigee, and stilllater to apogees at other locations overlying other continents andareas. Hence, that same satellite can later communicate with those otherareas. Therefore, this system allows a store-and-dump type system. Theinformation can be stored on board the satellite and laterre-transmitted when the satellite overlies those other areas. Thissystem also allows all the satellites in the array to communicate withthe other satellites in the constellation, through intersatellite links.This feature is desirable for real time communications.

This system has a number of other distinct advantages. Importantly, thesystem operation allows selecting specific geographic locations to bepreferentially covered; for example, continents can be followed by theconstellation to the exclusion of other areas, e.g. ocean areas betweenthe continents. The communication equipment on the continent alwayscommunicates with one satellite at apogee, although not always the samesatellite. From the point of view of the ground station, the satelliteappears to hover over the ground. This satellite system operatesvirtually like a geosynchronous satellite system. Importantly, thesesatellites according to the disclosed system orbit at about half thealtitude of the geo systems. A geo orbit orbits at 36,000 milesaltitude: the virtual geo satellite orbits at average altitudes of16-18,000 miles. Also, geo satellites require “apogee motors”, to boostthem from their original orbits into the final geo orbit. These apogeemotors can double the weight of the satellite. This yields acommunications system which costs less dollars per launch capabilitybecause of the reduced weight to boost and less size. Also, since thegeo satellites orbit at a higher altitude, they operate at a higherpower, and use a larger illuminating antenna, all other conditions onthe around being equal. These satellites also have a much larger overallsize. This size of the satellites increases as the square of thedistance. Therefore, the geo satellite needs to be at least twice aslarge and twice as powerful as a low altitude satellite. The powersupply conservation techniques of the disclosed system allow thesatellite to be made even smaller.

The system also provides satellites with very high elevation angles.Maximizing the elevation angle prevents interference with existingsatellites such as true geosynchronous satellites.

This is another feature of the disclosed system which allows thesesatellites to operate in ways which avoid any possibility ofinterference with the geo band. Another objective and important featureof the disclosed system is its ability to re-use satellite communicationchannels. Regulatory agencies such as the FCC allocate frequency bandsby allocating a specific frequency band for a specific purpose. The geosatellites, for example, receive an assignment of a frequency band.Thereafter, the regulatory agency will consider that other satelliteslocated in the same orbital position can not use this frequency becauseof possibility of interference. Hence, frequencies in adjacent bandswhich might interfere with that assigned band will not be allocated fornew satellite use. With the disclosed system, there is a large angularseparation between the geo-sats and those covered by the invention.Thus, the same frequencies ca be allocated anew. Another feature of thedisclosed system is the location of the earth stations and satellites ina way which prevents interference with the geo bands. Specifically, thedisclosed system defines embodiments using both inclined orbits andnon-inclined (equatorial) orbits. The inclined orbit embodiment of thedisclosed system only communicates with the ground stations when a linedrawn between the ground station and current position of the satellitewill not intersect any point within x.degree. of the ring ofgeosynchronous satellites, where x is the required separation betweenthe communication for geo satellites and the communication for thesatellites of the disclosed system. During other times, the equatorialcomponent of the communication is shut off. The satellite onlycommunicates when it is near apogee. During those times, the rotationalvelocity of the satellite approximates the rotational velocity of theearth, and hence the satellite tends to hang overhead relative to theearth.

For non-inclined (equatorial) orbits, the ground stations are placed ina position such that the communication does not intersect the ring ofequatorial orbits, by ensuring that satellite apogees are at loweraltitudes than apogees of geostationary satellites.

The system is controlled by on-board processor 280, which determines theposition in the orbit and the steering of the antenna from variousparameters. Processor 280 carries out the flowchart shown in FIG. 2 awhich will be described herein. The overall system is powered by powersupply 290 which supplies power to all of the various components andcircuitry which require such power. Power supply 290 includes a sourceof power, here shown as a solar array 292, and an energy storage elementhere shown as a battery array 294. Importantly, according to thedisclosed system, the solar array 292 is sized to provide only someamount of power less than that required to power the satellitecommunication. The amount by which the solar array can be less is calledherein the power ratio of the device. The power ratio depends on thekind of orbit that the satellite will have, and how long the satellitewill be transmitting during each elliptical orbit. The preferred powerratio is 0.5: this will power a satellite which is communicating halfthe time, and the other half the transmitter and receiver on board thesatellite is off and the solar array is providing power to chargebattery 294.

The flowchart of operation is shown in FIG. 2 a. Step 350 representscontrolling the antenna. This requires that the processor keep track ofthe satellite's position in the orbit. Step 352 determines if thesatellite is in a position in its orbit where it is active (transmittingand/or receiving). If so, flow passes to step 354 where power is drawnfrom power supply and the battery. If the satellite is not powered, thenpower is used to charge the battery at step 356.

The system also allows selective expansion of the communicationscoverage by adding additional satellites into additional ellipticalorbits.

The virtual geo satellite system of the disclosed system also enablescomplete communications coverage of the earth without requiring a groundnetwork. The same satellite services all different portions of the earthat different times of day. The coverage of the earth repeats over a 24hour period. A preferred embodiment receives information relayed fromthe ground, relays it to the earth area below it, then stores theinformation, and later reads back the stored information to retransmitthat same information to other areas of the earth. The system of thedisclosed system increases the satellite coverage at high densitygeographic locations using fewer satellites than was possible withprevious constellations by fixing the satellite apogee passages overgiven geographic regions defined by both longitude and latitude.

Integral values for mean motion of the satellites in the array ensuresthat the ground track repeats on a daily basis. The ground trackspreferably repeat each day so that the orbit apogee passes in the samelocation relative to the geographic target area. This system maximizesthe time of coverage and elevation angles for that pass.

Before describing the minimum satellite arrangement according to thedisclosed system, the nomenclature used herein to describe thecharacteristics of satellite orbits will be first described. The “meanmotion” is a value indicating the number of complete revolutions per daythat a satellite makes. If this number is an integer, then the number ofrevolutions each day is uniform. This means that the ground tracks ofthe satellites repeat each day: each ground track for each day overridesprevious tracks from the preceding day.

Mean motion (n) is conventionally defined as the hours in a day (24)divided by the hours that it takes a satellite to complete a singleorbit. For example, a satellite that completes an orbit every threehours (“a 3-hour satellite”) has a mean motion of 8.

The “elevation angle” .delta. is the angle from the observers horizon upto the satellite. A satellite on the horizon would have 0.degree.elevation while a satellite directly overhead would have 90.degree.elevation. Geo satellites orbit near the equator, and usually have a20-30.degree. elevation angle from points in the United States.

The “inclination” I is the angle between the orbital plane of thesatellite and the equatorial plane. Prograde orbit satellites orbit inthe same orbital sense (clockwise or counter-clockwise) as the earth.For prograde orbits, inclination lies between 0.degree. and 90.degree.Satellites in retrograde orbits rotate in the opposite orbital senserelative to the earth, so for retrograde orbits the inclination liesbetween 90.degree. and 180.degree.

The “critical inclination” for an elliptical orbit is the planarinclination that results in zero apsidal rotation rate. This results ina stable elliptical orbit whose apogee always stays at the same latitudein the same hemisphere. Two inclination values satisfy this condition:63.435.degree. for prograde orbits or its supplement 116.565.degree. forretrograde orbits.

The “ascending node” is the point on the equator where the satellitepasses from the southern hemisphere into the northern hemisphere. Theright ascension of the ascending node (“RAAN”) is the angle measuredeastward in the plane of the equator from a fixed inertial axis in space(the vernal equinox) to the ascending node.

The “argument of perigee” is a value that indicates the position whereorbital perigee occurs. When using equatorial orbits, 00 argument ofperigee is used for all the orbits. Inclined orbit arrays use non-zeroarguments of perigee. Arguments of perigee between 0.degree. and180.degree. locate the position of perigee in the northern hemisphereand hence concentrate the coverage in the southern hemisphere.Conversely, arguments of perigee between 180.degree. and 360.degree.locate the perigees to the southern hemisphere and hence concentrate thecoverage on the northern hemisphere.

An embodiment of the disclosed system evenly spaces the axes of theellipses. The spacing between RAANs is called “S” and calculated byS=360/n=120.degree.

The disclosed system positions the satellite coverage based on bothlongitude and latitude of the desired continental area to be covered bythe orbit. This is done, first, by synchronizing the orbit apogee topass over the targeted geographical region for each successivesatellite. We select a suitable value for the mean anomaly, which is afictitious angle relating to the elapsed time in orbit. 360.degree.represents the completion of the orbit. In this example, the meananomalies are also S=120.degree. apart.

This “mean anomaly” M relates the amount of time it takes the satelliteto rotate S.degree. around the earth (here 120.degree.). The meananomaly required for the 12-hour satellites to rotate to S.degree. is 8hours; two-thirds of a period. This corresponds roughly to the amount oftime the satellite remains in apogee. Taking the initial satellite nearapogee, therefore, (180.degree. mean anomaly) the next satellite shouldbe backed up by 240.degree. This means that after 8 hours that satellitewill be at 180.degree. Since 180.degree. minus 240.degree. is negative60.degree. which equals 3000, this is the value of mean anomaly M forsatellite number 2. This system is used to select values for theconstellation in a similar manner for each succeeding satellite.

Arrays with more satellites (“higher order arrays”) can also be madeusing the same rules as those discussed above. Successively largernumbers of satellites can be used to provide more coverage, moreoverlapping coverage, or smaller integral mean motion values. As thevalues of M get larger, the eccentricity of the ellipses become smaller.This is because the perigee altitude is fixed at about 500 km to avoidre-entry and decay into the earth's atmosphere; longer periods havehigher apogee altitudes greater supportable eccentricities.

FIG. 1A shows how the satellite ellipse is selected to have an angularrate in the plane of the equator, at apogee, which approximates theangular rate of the earth. The dotted line in FIG. 1A represents theangular rate of a geo satellite, and hence at this angular rate asatellite would approximate the angular speed of the earth. The ellipseis selected to have a semi-major axis length to set the minimum angularrate of the satellite at apogee. At apogee, the satellite angular rateshould approximate the rotational velocity of the earth. In reality,this rotational velocity will be either a little faster or a littleslower than the earth. At this time, therefore, the satellite appears tohang relative to the earth. All elliptical orbits, including whosedescribed herein, are also subject to effects of long-termperturbations. If effects of these long term perturbations are notcompensated, this could cause continental coverage to drift with thepassage of time.

These perturbation effects are mainly effects from the Earth's J2rotation harmonic. The earth is not a perfect sphere; it actually bulgesat the equator. This causes gravitational effects on objects which orbitthe earth. For posigrade orbits (i>90.degree.) the line of nodes willregress. For inclinations greater than critical (63.4.degree.>i>116.6),the line between the perigee and apogee (line of apsides) will regress;for other inclinations, I<63.4.degree. or I>116.6, the line of apsideswill progress. Exactly at the critical angles I=63.4 or I=116.6, theline of apsides will remain stable a very desirable feature inmaintaining apogee at a certain latitude. In the equatorial plane, thecombined effect of these two major perturbations cause the apogee toadvance or move counter-clockwise from the sense of looking down fromthe celestial north pole. All of the satellites in a given array designwould be affected similarly. Fortunately, this effect could becompensated by slightly increasing the period of each satellite in thearray by an amount which offsets the J2 perturbation. This affects thesystem by causing a point on the earth to take a slightly longer time toreach the satellite's next apogee arrival point. This effect iscompensated by slightly increasing the satellite's period. The advanceof perigee is suppressed by setting the inclination at one of thecritical values.

A first embodiment of the invention uses N=3 satellites, where N is thetotal number of satellites, preferably in the equatorial plane, to coverN−1=2 continents. The rules for spacing and phasing the satellites willbe given in the general form that can be used later for more complicatedconstellations or arrays. The mean motion integer sets the minimumnumber of satellites in the array and n.sub.c the number of continentsthat are followed. Here n.sub.c=2 provides a satellite period equal to12 sidereal hours. N (the minimum number of elliptic satellites in thearray) is determined by using the relationship N=n.sub.c+1. Thus, N=3.This is the minimum number of satellites that need to be in the array;we can also set the number of satellites in the array N to be anyinteger greater than n+1.

The apogee passage is synchronized over the targeted geographicalregion, for each successive satellite, moving counterclockwise as viewedfrom the celestial North Pole. This is accomplished by selecting asuitable value for the mean anomaly.

Refinements: Additional features augmenting the usefulness of the abovesimpler version include:

1) Inclining the elliptical orbital planes at the critical inclinationangles (63.435 or 116.535.degree.), with phasing to maintain a singlerepeating ground track. The single repeating ground track for thesimplified non-inclined example above is simply the line of the equator.

2) Taking advantage of the higher apogees in allowing more directcross-linking between satellites than with present low-altitude circulararrays. Usually, a single cross-link suffices, even when the longitudedifference between end points is 180.degree. (on the opposite side ofthe earth).

3) Placement of apogees over a selected latitude and longitude foroptimal coverage of a potential market area. This is done through properselection of all the orbital parameters, with particular attention givento selection of argument of perigee, .omega.

First Embodiment

The orbits of the disclosed system are shown in FIG. 1. The satellite100 is shown in an elliptical orbit 102 around the earth. Thecommunication equipment on the satellite 100 communicates with earthground station 104, and also beams the information to earth groundstation 106. Satellite 110 is shown in a separate independent ellipticalorbit communicating with ground stations 112 and 114 on the earth. Notealso that the satellite 100 can communicate directly to the satellite110 via communication link 120.

The preferred characteristics of these orbits are described in Table I.

TABLE I Satellite No. P1 P2 P3 Semi-Major 26553.98 km 26553.98 km26553.98 km Axis, a=Inclination, I=0 deg 0 deg 0 deg Arg. Perigee, 270deg 270 deg 270 deg w=Eccentricity, 0.51 0.51 0.51 e=Rt. Ascension, 0deg 120 deg 240 deg RAAN=Mean 180 deg 300 deg 60 deg Anomaly,MA=Satellite 100 also includes store and dump hardware thereon asdescribed herein. This allows the satellite to obtain programinformation so that later in its orbit, when at the position 130, it cansend its same information to ground station 132.

A detailed block diagram of the electronics in the satellite is shown inFIG. 2. This block diagram shows elements which carry out communicationbetween the ground station 104, the satellite 100, and the remote userstation 106. The inter-satellite links 120 are shown from the satellite100 to the satellite 110. The video input to be distributed is receivedas video input 200, and input to a video coder 202 which producesdigital coded video information. This digital coded video is multiplexedwith a number of other channels of video information by videomultiplexer 204. The resultant multiplexed video 206 is modulated andappropriately coded by element 208 and then up-converted by transmitterelement 210. The up-converted signal is transmitted in the Ku band, ataround 14 GHz, by antenna 212. Antenna 212 is pointed at the satellite100 and received by the satellite's receive phased array antenna 214.Antenna 212 is controlled by pointing servos 213.

The received signal is detected by receiver 216, from which it is inputto multiplexer 218. Multiplexer 218 also receives information from theinter-satellite transponders 240.

The output of multiplexer 218 feeds the direct transponders 250, whichthrough a power amplifier 252 and multiplexer 254 feeds beam former 256.Beam former 256 drives a transmit, steerable phased-array antenna 260which transmits a signal in a current geo frequency band to antenna 262in the remote user terminal 106. This signal preferably uses the samefrequency that is used by current geo satellites. The phased arrayantenna is steered by an on-board computer which follows a pre-set andrepeating path, or from the ground. This information is received byreceiver 264, demodulated at 266, and decoded at 268 to produce thevideo output 270.

The satellite includes another input to the multiplexer from thesteerable antenna, via the intersatellite link 120 and receiver 240.Transmit information for the the intersatellite link is multiplexed at242 and amplified at 246 prior to being multiplexed.

Output 222 of input multiplexer represents a storage output. Thesatellite electronics include the capability to store one hour of TVprogram information. The TV channels typically produce information atthe rate of 6 megabytes per second. The channels are typically digitallymultiplexed to produce information on 4-6 channels at a time. Therefore,the disclosed system preferably uses 22 gigabytes of storage to storeover 1 hour of information at about 4.7 megabytes per second. Theinformation stored will be broadcast over the next continent. Thestorage unit 224, accordingly, is a wide SCSI-2 device capable ofreceiving 4.7 megabytes per second and storing 22 GB.

Upon appropriate satellite command, the output of the storage unit ismodulated and up-converted at 226.

This basic system shown in FIG. 2 can be used in one of the preferredsatellite arrays of the disclosed system. These arrays will be discussedherein with reference to the accompanying drawings which show thecharacteristics of these satellite arrays.

This first embodiment uses a simplified 12-hour equatorial planesatellite array n=2, N=3. The mean motion n of 2 means that eachsatellite completes an orbit around the earth twice per day.

An important enhancement of an N=3 case is obtained by modifying thecharacteristics of the orbits so that the satellites coalesce over thecovered areas at the moments when satellite coverage changes. The termcoalesce as used herein means that as one satellite moves out of rangeof the ground tracking, the next satellite moves into range at that sameposition. In fact, the two satellites come very close to one another atthat point—within 1.degree. from the view of the satellite. Thissimplifies the ground tracking, since the switchover between satellitesdoes not require much antenna movement.

FIGS. 4A-4F show the basic three-satellite “rosette” formed by the threeelliptical orbits. The earth 300 is located at one of the foci of eachof the three ellipses of the respective satellites. Satellite 302communicates with point 304 on the earth. Satellite 302 orbits the earthin ellipse 306. The satellites 1, 2 and 3 respectively have ascendingnodes of 0, 120 and 240, and respectively have mean anomalies of 180,300, and 60.

Similarly, satellite 310 orbits the earth in ellipse 312, and satellite320 orbits the earth in ellipse 322. Satellites 310 and 320 are both ina position to provide coverage to the second covered continent area 314.Note that satellites 310 and 320 are in their coalesced position—theyare very close positionally, to one another. Satellite 320 is movingaway from apogee while satellite 310 is moving toward apogee. Thetracking antenna is hence commanded to switch between tracked satellitesa the time when satellites 310 and 320 are positionally very close, buthaving adequate angular separation to avoid self-interference. Accordingto the disclosed system, this switchover occurs when the satellites arewithin 5.degree. of each other.

The satellites all orbit in a counter-clockwise direction relative tothe sense shown in FIG. 4. The earth also orbits in thecounter-clockwise direction. The semi-major axes of the ellipses in FIG.4 are shown as axes 308, 314, and 316, respectively.

In order to describe these orbits, first the characteristics of anellipse will be described. FIG. 3 shows ellipse 400, having a focus 402.The satellite orbits along the path of the ellipse 400, with the centerof the earth being at the focus position 402 (“the occupied focus”).

The apogee 404 and the perigee 406 of the orbits are defined by thepoints on the ellipse which are farthest from and closest to the focusof the ellipse, respectively. The amount of difference between thesedistances define the eccentricity of the ellipse. The semi-major axis408 is defined as half of the long axis of the ellipse. This semi-majoraxis runs through the two foci of the ellipse, to split the ellipse intotwo halves. The two lengths along the semi-major axis, from one edge ofthe ellipse to the occupied focus of the ellipse are called the “radiusof perigee” and the “radius of apogee”; the latter being the longer.

As the eccentricity of an ellipse approaches zero, the ellipse becomesless elliptical, eventually approaching a circle (e=0) when theeccentricity is zero. The semi-major axis of a circle is the radius ofthe circle.

The characteristics of the ellipse/object in elliptical orbit arecalculated as follows.

The apogee, r.sub.a=a.multidot.(1+ECC).

Perigee r.sub.p=a.multidot.(1−ECC).

A more eccentric ellipse (higher value of eccentricity ECC) has agreater difference between the values P and R. Hence, such an ellipse isless like a circle. The characteristics of the ellipse are thereforedetermined as a function of its eccentricity.

The position of a satellite in orbit follows Kepler's laws of motionwhich states that the orbiting element will sweep out equal areas of theorbit in equal times. This results in the satellite moving very rapidlywhen it is at an approaching perigee, but very slowly when it reachesapogee. For a twelve hour elliptical orbit, therefore, it can be seenthat the satellite will spend most of its time near apogee. The numberson the ellipse of FIG. 3 represent time indications of hours passed in a12 hour orbit, e.g., they indicate the number of hours since zero thathave elapsed in a 12 hour orbit.

The preferred ellipse for the 3-satellite elliptical orbit has aneccentricity of about 0.51. This value best allows the satellites tocoalesce.

The earth rotates once in every 24 hour period, and hence takes eighthours to rotate between the major axes of the three equally spacedellipses (120.degree. spacing); FIG. 4A shows the point to be covered304 is initially pointing directly towards satellite 302 which is atapogee at time 0:00. As time passes, both the satellite 302 and theearth will rotate.

As time passes, the satellites move from the position shown in FIG. 4A.FIG. 4B shows the position one hour later at time 1:00. Satellite P1 hasmoved away from apogee, although it has moved relatively little.Satellite P2, on the other hand, is now moving much more rapidly at thistime, since it is approaching perigee, while P3 is still near the apogeeposition.

An observer on or near the equator sees the nearest satellite appear toclimb in altitude from almost directly overhead, towards apogee, all thewhile staying almost directly overhead at an elevation angle of 80-90.The satellite is actually rotating more slowly than the earth duringthis time: it is appearing to move from east to west, rather than westto east as most low or medium altitude satellites move in the sky.

FIG. 4C shows a view of the satellites one hour later at time 2:00. Thetracked locations 304 and 314 each still view a satellite near itsapogee position. Satellite P3 continues to move towards apogee and henceappears to hang overhead. P1 is still around apogee and thus alsoappears to hover.

FIG. 4D shows yet another hour later at time 3:00. P3 is still atapogee, but P1 is approaching perigee. Notice that P2 is coming out ofperigee and approaching the coalescence point at which P1 and P3 willcross paths. That crossing of paths is shown in FIG. 4E, time 4:00, whenP1 and P2 have coalesced in their positions at the time when point 304switches over between coverage by satellite P1 and P2. At that time, thesatellites are within 1.degree. of one another as viewed from theground.

The above has described the satellite P1 moving from directly overheadthe point to be covered, to the point where satellite P1 no longercovers the point to be covered. Therefore, the satellite is transmittingfor eight of the twelve hours of its orbit; ⅔ of the time.

This cycle repeats. As the satellites continue to orbit, differentsatellites take similar positions to those shown in FIGS. 4A-4E. FIG. 4Fshows the cycle starting to repeat with satellite P2 moving towardapogee, satellite P1 moving toward perigee, and P3 hovering relative tothe earth near its apogee.

FIGS. 4A-4F demonstrate the important features recognized by theinventors of the disclosed system, whereby the satellites spend most oftheir time at apogee. At the highest points of apogee, the velocity ofthe satellite very nearly matches that of the earth, and so thesatellite appears to hang overhead. The satellite is preferably trackedwhile its angular velocity differs from the earth's angular velocity by20% or less.

Importantly, the covered areas on the earth always see either asatellite directly overhead or two satellites which are very nearlydirectly overhead. FIGS. 4A-4F show how this system actually appears tothe communications point 304 to be virtually geosynchronous. Thecommunications point communicates with different satellites at differenttimes in the satellite orbit. The communications point is alwayscommunicating with one satellite.

The satellites follow repeating ground tracks, since the cycle ofsatellite movement shown in FIGS. 4A-4F continually repeats.Importantly, this allows the ground tracking antenna 212 to continuallyfollow the same path, starting at a beginning point, tracking thesatellite, and ending at the coalesce point. After the satellitescoalesce as shown in FIG. 4A, the antenna begins its tracking cycle. Theinventors of the disclosed system have optimized this system forpreventing interference with geo satellites.

Specifically, consider FIG. 4G which shows a multiplicity of satellitesin inclined elliptical orbits. The disclosed system preferably operatesto monitor satellites at and near their apogee positions. The satellitesnear perigee are moving too rapidly, and hence are not tracked. Moregenerally, the system of the disclosed system operates such that thesatellites are only being used at certain times during their orbits. Inthis embodiment, those certain times are when the satellites are atapogee. Non geosynchronous circular arrays are commonly used at present;they are actually much less efficient, since with zero eccentricity theyspend a significantly greater time on the side of the earth away fromthe populated continents. The arrays of the disclosed system, on theother hand, spend most of the time at or near apogee over the populatedcontinents of interest, and a relatively small time (at high angularvelocities) passing through perigee in regions of no commercialinterest.

The satellites are only used when their geometry is such that there isno possibility of the line of sight between the ground station and thesatellite interfering with the geosynchronous band of satellites. Thisallows the satellite communication to take place on the samecommunication frequency band normally assigned to geosynchronoussatellites.

Moreover, the disclosed system teaches that when the satellites are notcommunicating, either because the satellites are no longer at theirtracked apogee portion and/or when the satellites are in a region wherethey might interfere with geosynchronous satellites, the maintransmission is turned off. During this time, the power supply is usedto charge the battery. This means that the power supply can be madesmaller by some factor related to the duty cycle of the satellite.

Another consideration is since the satellites only communicate whilenear apogee, they are never eclipsed by the earth. The satellites canalways receive sunlight for solar operation while transmitting andreceiving.

For example, FIG. 4G shows satellites in orbit. In the example given inFIG. 4G, the satellites are only tracked when they are in the positionof the orbit above the line 450. The only possibility of interferencewith geo satellites comes when the tracking beam is within 10.degree. to30.degree. of the geo band. So long as an angular separation greaterthan this amount is maintained, there can be no interference. Therefore,the disclosed system allows re-using the frequency bands which areusually assigned to geosynchronous satellites in a position whereinterference with the existing satellites can not occur.

The same rules are used to construct higher order arrays withsuccessively larger integer mean motions and hence shorter periods.These arrays require a larger number of satellites, but provide somewhatbetter coverage of the earth. Since more satellites are used in thesehigher order arrays, each satellite need spend a lesser amount of itstime at apogee. This allows orbits to be formed wherein the values ofeccentricity are allowed to become smaller as the mean motion increases.The ultimate limit is atmospheric drag, which limits perigee altitudesto about 500 kilometers. This would correspond to a 1500 kilometerapogee elliptical orbit with a resulting eccentricity of(r.sub.a−r.sub.p)/(r.sub.a+r.sub.p) which is approximately 0.067. Thisdescribed orbit is not practical since its period is about 1 hour and 45minutes which is not an integral value for the mean motion. The nextnearest value for mean motion would be n=14. The n=14 orbit, however,would be so slightly elliptic that it would not offer much advantageover the circular orbits.

Practically, those arrays having mean motions of 3, 4, 5, 6, 7 and 8 aremost preferred according to the disclosed system. The most preferredorbits according to this invention include the three-satellite orbits,the four-satellite orbits, and the five-satellite orbits. A particularlyadvantageous embodiment uses two arrays of five satellite orbits.

As discussed above, all of these orbits include long-term perturbationswhich would, if not compensated, cause the desired continental coverageto drift off with the passage of time. The two major perturbationeffects are due to the earth's J.sub.2 harmonic; and include:

Regression of the line of nodes (for posigrade orbits), and Advance ofperigee.

For inclined orbits, the advance of perigee can be suppressed by settingthe inclination, i, at either 63.435 or 116.565.degree.

The combined effect of these two major perturbations in the equatorialplane, due to the J.sub.2 harmonic term has the net effect of causingthe apogee to advance in a counter-clockwise direction looking down fromthe celestial North Pole.). All the satellites in a given array designwould be affected alike. Fortunately, this effect can be compensated byincreasing slightly the period of each satellite in the array in a waysuch that the earth takes a slightly longer time to reach the nextsatellite's apogee arrival point. This is compensated by adding thisextra time to the satellites' periods. The exact amount will vary, andis a function of a number of variables, including the orbital periods,inclinations, and eccentricities. For inclined elliptic orbits (atcritical inclination angles), there will be no rotation of perigee ineither direction. However, there will be a regression of the line ofnodes which must be compensated by a small adjustment in orbital period.This will cause the plane of the orbit to rotate clockwise in the senselooking down from the North Pole. If that happens, the satellite wouldpass over a selected meridian at a slightly earlier time each day (oreach repeat cycle), unless we adjust the period of the satellite. Inthis case, we would shorten the period of the satellite, whicheffectively ‘stretches’ out the trajectory ground trace and causes theground track to repeat exactly over the life of the satellite.

As described above, third order effects due to tesseral terms may needto be compensated by small orbit maintenance maneuvers using minusculeamount of fuel.

The preferred four-satellite array is shown in FIGS. 5A-5E. This arrayshows four satellites used to track three continents. These satellitesorbit in elliptical orbits having an eccentricity of 0.6. FIGS. 5B and5D show the satellite coalescing which occurs according to thisembodiment.

FIG. 6 shows an overall view of the 10 satellite array; and FIGS. 7A-7Eshow the ground tracks for a satellite array with 5 satellites having aperiod, T, equal to 6 hours. This array is preferably used with two setsof five satellites, yielding a ten-satellite, six hour constellation.

The preferred communications system uses a ten satellite system, eachhaving six hour orbits, and each optimized for users in the Washington,D.C. area. This still, however, provides coverage throughout the rest ofthe continental United States, and the entire northern hemisphere aswell as that part of the southern hemisphere down to about 10 deg Southlatitude.

The system uses ten equally-spaced prograde satellite orbit planes. Allsatellite orbits are at the ‘critical’ inclination angle of63.435.degree. to prevent rotation of the line of apsides.

The ground track is adjusted so as to pass directly over Washington,D.C. by adjusting the right ascensions of all the orbits whilemaintaining their equal spacing. The argument of perigee is adjusted toobtain apogees over or nearly over the targeted latitude and longitude.

FIG. 6 shows an overview of the orbital constellation. It can readily beseen that the satellites favor the Northern Hemisphere by spending moretime, and reaching a higher altitude in the Northern Hemisphere. FIG. 6shows a snapshot of time at 0:00 hours, and it should be seen that allsatellites except for satellites P5 and P1 are over the NorthernHemisphere at that time.

FIGS. 7A-7G show a Cartesian, or Mercator, plot of the world showing therepeating ground tracks. The satellite array has a repeating groundtrack that repeats every 24 hours. The satellites appear to ‘hover’ ordwell along four equally-spaced meridians, one of which is at thelongitude of Washington, D.C.; the others being spaced at 90.degree.intervals from Washington.

FIG. 8 shows the minimum elevation angle to the highest satellite overWashington, D.C., as a function of time. Every 24 hour period has tenelevation angle peaks of satellites on a descending (from northproceeding towards the equator) at or near the observer's zenith (90deg). The lower, sharper peaks in the figure represent other satelliteson ascending passes; they are at lower altitudes and thus going faster.These ascending satellites are not actively transmitting to users on theground at the times when they are on ascending passes.

The preferred system uses a total of ten (10) satellites incritically-inclined (i=63.4 deg) 6-hour orbits, phased and oriented toprovide optimal earth coverage. As will be seen, this geometry alsoprovides a very high elevation angle, and hence avoids interference withthe existing geo communications satellite band. The preferred orbitshave apogee and per-gee altitudes of 20074 and 654 kilometers,respectively.

From a user's viewpoint, the satellites are accessed sequentially atnominal 2 hour and 24 minute intervals at exactly the same point in thenorthwestern sky (the ‘start point’ of the tracking segment), and aretracked in a roughly northwest to southeast trajectory to a point in thesky well short of intersecting the geo band of satellites. Thesatellites remain at apogee during the time while they are being trackedfrom the ground. Hence, these satellites are only tracked, andcommunicated with, while their velocity closely matches the velocity ofthe earth. When the satellites begin to approach the perigee stage, andhence their velocity increases relative to the earth's rotation todiffer therefrom by more than 25%, for example, they are no longer beingtracked by the communication equipment on the earth. At this end pointof the tracking segment, the ground communications antenna is directedback to tracking its start point to repeat the sequence as thenext-appearing satellite is acquired. Tracking along the active arcsegment is accomplished at less than 2 deg/min. For the present array,this results in every ground communications antenna effecting tenswitchovers per day. As explained above with reference to FIG. 1, thesteering operation of the disclosed system preferably uses phased arraysteering of the antenna. However, more-conventional antenna steering isalso contemplated.

Importantly, the trajectory segments appear exactly the same to the userfor every satellite, since the azimuth-elevation trace is repeated foreach satellite. This system defines significant advantages. Itsoperating altitudes are half that of existing geo systems. This greatlyreduces link margins and emitted power requirements for the satellites.

Apogees are placed on the meridians of longitude of theheavily-populated areas for which the constellation is optimized. Apogeepoints may also be adjusted to approximate the targeted area latitudesas well. The satellite tracking arcs over the targeted areas remainroughly overhead (within 30-40.degree. of zenith), with slow angularmovement during periods when the satellite is active. The trajectoriesfor mid-latitude (20-50.degree. North latitude) observers locateddirectly under the apogee points in the high-population targeted areasare approximately north-south oriented.

All ten ground tracks are identical, and only the satellite that iscurrently covering the repeating ground tracks charge. The repeat cycleis 24 hours. Since the satellites move from one geographic area toanother, information once transmitted can be re-broadcast at anotherlocation.

The Mercator plot of FIGS. 7A-7E show that the entire system actuallyfollows one ground track, repeating after 24 hours. It actually ‘foldsover’ from the left edge of the world map to the right edge, giving itthe appearance of multiple traces.

Table II gives the orbital parameters, or ephemerides, of the entirearray of ten satellites:

TABLE II SYSTEM ORBITAL PARAMETERS Sat RMN MA # a(km) i(deg) e,(ecc.)w,(deg) (deg) (deg) 1 16742 63.435 0.58 315 0 0 2 16742 63.435 0.58 315072 072 3 16742 63.435 0.58 315 144 144 4 16742 63.435 0.58 315 216 2165 16742 63.435 0.58 315 288 288 6 16742 63.435 0.58 315 180 0 7 1674263.435 0.58 315 252 072 8 16742 63.435 0.58 315 324 144 9 16742 63.4350.58 315 036 216 10 16742 63.435 0.58 315 108 288

Some adjustments will be required to account for long term orbitalperturbations as described above. This adjustment is common insatellites requiring precise repeat cycles such as Topex-Poseidon, orthe Canadian Radarsat.

Similar views to those from the above can be drawn for the preferredten-satellite array. An important point of the ten-satellite array,moreover, is that there is good inter-satellite connectivity forcross-linking.

FIG. 7A shows the position of the satellites at time 00:00. Compare thiswith FIG. 7B, which shows the same satellites twenty-four minutes later.The satellite P4, which is substantially over Washington, D.C., hasmoved very little, albeit P5 will be picking up speed as it approachesperigee. P4 appears to hang over Washington, D.C., since it is near theapogee portion of its orbit and its velocity very closely matches thevelocity of the earth.

In contrast, during the same short period of time, the satellite P1, atperigee, has moved very quickly and very far along its orbit. Similarly,satellite P8 (over Europe), P5 (over Southern Africa) and P9 have movedvery little. Twenty-four minutes later, FIG. 7C shows that satellite P4has started to move away from the United States, but satellite P7 is nowin place, very close to its apogee. This is evident from its positiontwenty-four minutes after that, shown in FIG. 7D, where satellite P7 hasmoved only very little, and is still well-covering the United States. Attime 1:36 shown in FIG. 7E, the satellite P7 is over Washington, D.C.

The satellite P7 is still over Washington D.C. at time 2:00 hours, shownin FIG. 7F. The satellite starts to move at time 2:24, shown in FIG. 7G.

The disclosed system intends that the satellites be used forcommunication during only some part of the time while they are in orbit.During other times in orbits, the satellites are not being used forcommunication, but instead are charging their energy storage. Thisfeature of the invention has been described above, but will be describedin more detail herein with reference to FIGS. 2A, 4G and 4H.

FIG. 4G shows a view of the earth from, for example, the view of thesatellite from the sun. This figure shows all of the satellite orbits,and their elliptical orbital paths. The geosynchronous satellites are inequatorial planes shown as the geo ring 800. Communications equipment onthe earth communicates with this geo ring 800. Moreover, sometimes thegeo satellites are perturbed by the earth's oblateness, henceeffectively forming orbits which are slightly inclined. The geo ringsshould therefore be considered at occupying a 5′ position borderingtheir nominal position.

Ground communications equipment on the earth communicates with this georing. The cone of communications to the geo ring is shown as 802.

When the ground communication equipment on the earth communicates withthe satellites P1-P5, it should be seen that they are aimed at aposition of the sky, 804, which is completely separated from the georing 802. According to the disclosed system, a distance is maintainedbetween the satellites and the geo ring 800. The angular separation ethe minimum acceptable angular separation which can ensure nointerference between the geo ring and the satellites of the disclosedsystem. An embodiment uses an angular separation of 30.degree., which isan amount which will obviate any possibility of interference problem.More generally, however, any angular separation greater than 15.degree.would be acceptable.

Taking the satellite P3 as an example, therefore, the satellite can onlybe used according to the disclosed system when it is in its orbitbetween the points labelled 808 and 810. However, the virtual geo systemwhich is preferably used according to the disclosed system uses thesesatellites during even less of their orbit, only between the points 812and 814. When the satellite is in the other positions of its orbit, thesatellite is not consuming power or transmitting. Therefore, thisprevents any possibility of interference with the geo satellite systems.

The operation of the equatorial satellites is similar. The equatorialsatellite array is shown in FIG. 4 h. The equatorial satellite is shownas satellite ring 850. If the ground station is on the equator, shown asground station 852, then it would, at least at some times, interferewith satellites in the geo ring shown as 854. However, if the groundstation is separated from the equator by at least 30.degree., such asshown as position 856, then at least part of the satellite ring has nochance of interference with the ring 854. Therefore, the satellitecalculates geometries such as to obviate interference with the satellitering. Therefore, more generally, the disclosed system operates as shownin FIG. 2 a. The antenna is controlled at step 350, and from the antennacontrol the position of the satellite relative to geo are determined atstep 870. This can be determined, for example, from the pointing angleof the antenna. Step 872 determines if there is any possibility ofinterference between the two. This is determined from a numericaldifference between the pointing angle and the position of the geo ring.If there is any possibility of interference, control passes to step 874where the satellite communications is disabled. If interference is notpossible at step 872, then the satellite is enabled at step 874. Anenabled satellite can be, but is not necessarily, turned on. Forexample, in the virtual geo embodiments, the enabled satellite will bemaintained in the “off” position during some of the time when it isenabled. Therefore, step 352 determines if the satellite is powered.This may be determined from the repeating ground track, or otherinformation. If the satellite is not powered at step 352, the battery ischarged at step 356. If the satellite is powered, then power is drawnfrom both the supply and the battery at step 354.

Second Embodiment

Another embodiment, also referred to herein as the “VIRGO” embodiment,uses satellite sub-constellations with prograde elliptical orbits ofapproximately 8 hour periods. Each of the satellites within asub-constellation has the same ground tracks as the other satelliteswithin the subconstellation, or repeating ground tracks.

Each sub-constellation includes several satellites in each of theindividual ground tracks. The satellites are spaced such that as onesatellite leaves a service area, another satellite replaces it in thesame ground track.

As will be established herein, each satellite is in communication withthe ground station during a portion of the trajectory where thesatellite is at or near its apogee. During this time, the relativemotion of the satellite, i.e. the perceived motion of the satelliterelative to the Earth, is slow. The satellite travels through arelatively small angular arc, e.g., 40%, during its active phase.

As the one satellite departs from its active phase in the descendingdirection, the ground user can switch to the next-appearing satellite inthe ascending portion of the active phase of this next satellite.Continuity of coverage is thus provided by this switch-over.

During its active phase, each satellite is virtually geostationary. Thatmeans that it appears relatively stationary to a user on the earth.

The concept behind the virtual geostationary orbit can be illustratedwith analogy to the walking juggler. A juggler's clubs cluster togetherand move very slowly at the highest point in their trajectories. At thelow point of the trajectories, the juggler is catching and transferringthe clubs hand-to-hand rapidly. At the high point of the trajectory,however, the clubs move much slower.

The satellites in a virtually geostationary constellation areintentionally placed in stable elliptical orbits with their apogees overthe intended users. Like the juggler, these portions rise over theservice area and appear to hang there. Additionally, each satellite isactive for only a predetermined portion of its orbiting time, closest toits apogee portion. The satellites are spaced such that when onesatellite in the subconstellation reaches its inactive portion, anothersatellite in the subconstellation becomes active. Hence, the satellitesare spaced such that one ascending satellite replaces another descendingsatellite leaving the service area.

Since the satellites are in 8 hour orbits, each satellite peaks threetimes in each 24-hour day. Each of the peaks is located to follow apopulated region. Using a Northern Hemisphere apogee orbit as anexample, each satellite ascends, reaches its turn on point and beginsoperating, goes through its peak (“apogee”) and then descends. Thesatellite eventually reaches its turn off point. The satellite is thenreplaced, after its time of “hanging”, by the next satellite in thearray. The first satellite then falls rapidly into the SouthernHemisphere and quickly rises into the next Northern Hemisphere peak.Each satellite's peak is placed over one of the three NorthernHemisphere Continental masses each day. In order to provide coverage tocountries in the Southern Hemisphere, the embodiment employs anothergrouping of 5 satellites having their apogees in the SouthernHemisphere.

Each of the subconstellations is a mean motion 3 array. Each of thesatellite peaks is separated from other satellite peaks by 120.degree.of longitude (360.degree./3).

The longitudes selected for apogee placement of this array are79.degree. W, 41.degree. E, and 161.degree. E longitude. These fivesatellites serve the populated areas of South America, South Africa,Australia and New Zealand. Each satellite, in a single day, appears atapogee three times. This requires three satellites out of a total offive to be active at any time. Overall, each satellite must then beactive ⅗ of the time over a full day, or 14.4 hours. Since thisrepresents one day's total active time, and the satellite has beenactive over three geographic region, each region will be covered by asingle satellite for 4.8 hours. In other words, each 8-hour satelliteperiod, the satellite will be active for a 4.8 hour period—or 2.4 hourson either side of the apogee.

The satellites in this array have a duty cycle of 60%; that is, they areactively communicating 60% of the time. Their on/off switching timesoccur 2.4 hours on either side of the apogee. This corresponds to alatitude of 46.degree., and an altitude of 18044 km. The active phasefor each satellite occurs at latitudes greater than 46.degree. andaltitudes greater than 18044 (up to and including apogee at 27288 km).The satellites remain well clear of the GEO band, while active, so thereis no possibility for electronic interference with GEO communicationsatellites.

Because of the operating features discussed above, VIRGO satellitesoperate only when the satellites are at least 40.degree. separated fromthe line of sight of geo satellites. Hence, existing Ku and C frequencyequipment can be used without interfering with other communication.

The elliptical planes in the two Northern Hemisphere sub-constellationsare inclined at 63.40 with respect to the plane of the equator. Thismeans that the apogees will always appear to be roughly at 63.40 Northlatitude.

The two 5-satellite sub-constellations are called Aurora 1 and Aurora 2.These are used to provide continuous coverage of this type. The thirdsubconstellation is called Australis. Two or three spare satellites areplaced into “parking” orbits where they can be boosted into differentorbits if necessary.

The VIRGO™ orbital characteristics are as follows.

TABLE 1 VIRGO™ ORBITAL CHARACTERISTICS Aurora I™ Aurora II™ Australis™Spare Sats n=1-5 Sats n=1-5 Sats n=1-5 Satellites Semimajor 20281 2028120281 7285 Axis Eccentricity 0.66 0.66 0.66 0.05346 Inclination 63.43563.435 63.435 63.435 Right 341.5 255.3 52.2 0 Ascension of 53.5 327.3124.5 the Ascending 125.5 39.3 196.5 180 Node 197.5 111.3 268.5 269.5183.3 340.5 30 Argument of 270 270 90 270 Perigee 270 270 90 270 270 27090 90 270 270 90 270 270 90 Mean Anomaly 0 108.2 0 0 144 252.2 144 0 28836.2 288 0 72 182.2 72 216 324.2 216

This apogee of these VIRGO™ satellites is at 27,300 kilometers. This isapproximately three-quarters the altitude of geostationary satellites.This lower altitude provides less propagation delay to orbit.

The ground tracks of this embodiment are shown in FIG. 9. These producethe following locations of VIRGO™ active arcs.

TABLE 2 LOCATIONS OF THE VIRGO™ ACTIVE ARCS (Sub-Satellite Longitudes inDegrees East) AURORA I™ AURORA II™ NORTHERN NORTHERN AUSTRALIS™HEMISPHERE HEMISPHERE SOUTHERN HEMISPHERE 8-53 78-123 19-64 EuropeIndia-China Africa 128-173 198-243 139-184 Japan Alaska-HawaiiAustralia-NZ 248-293 318-3 259-304 Con.US N. Atlantic South America

Further information on the ground track is shown in the following.

TABLE 3 VIRTUAL-GEO ORBITAL ELEMENTS, PREFERRED EMBODIMENT AllSatellites: Semi-Major Axis (a)=20381 km; Eccentricity, e,=0.66;Inclination, I,=63.435.degree. Sat. Ground Track No. RAAN omega. MA #1(West. US) VG1a 350 270 0 #1 (West. US) VG2a 62 270 144 #1 (West. US)VG3a 134 270 288 #1 (West. US) VG4a 206 270 72 #1 (West. US) VG5a 278270 216 #2 (East. US) VG1b 263.8 270 108.2 #2 (East. US) VG2b 335.8 270252.2 #2 (East. US) VG3b 47.8 270 36.2 #2 (East. US) VG4b 119.8 270180.2 #2 (East. US) VG5b 191.8 270 324.2 #3 (S.A., Australia) VG1 c 6190 0 #3 (S.A., Australia) VG2c 133 90 144 #3 (S.A., Australia) VG3c 20590 288 #3 (S.A., Australia) VG4c 277 90 72 #3 (S.A., Australia) VG5c 34990 216

Although only a few embodiments have been described in detail above,other embodiments are contemplated by the inventor and are intended tobe encompassed within the following claims. In addition, othermodifications are contemplated and are also intended to be covered.

1. A satellite communications system, comprising: a ground station,including communications equipment and an antenna, located at a positionon the earth; a plurality of satellites in orbits around the earthhaving apogees and perigees, each of the satellites havingcommunications equipment thereon configured to communicate with theground station only during a predetermined portion of the satellite'sorbit proximate to apogee, the orbits of the plurality of satellitesbeing configured to form at least two ground tracks on the earthdisplaced from each other longitudinally, each of the ground tracksrepeating daily and having a number of active arcs, each active arccorresponding to the portion of the orbit of each satellite during whichthe communications equipment on the satellite is enabled to communicatewith the ground station, the orbits of the plurality of satellites beingfurther configured such that at all times there are at least two of thesatellites in each of the active arcs and such that at all times each ofthe satellites in any one of the active arcs is separated by at least apredetermined angle, as observed from the ground station, from eachother satellite in the same active arc and from any satellite in anyother active arc.
 2. A system according to claim 1, wherein the orbit ofeach of the plurality of satellites has a mean motion that is one of 2,3 and
 4. 3. A system according to claim 1, wherein the orbits of each ofthe plurality of satellites is inclined at critical inclination.
 4. Asystem according to claim 1, wherein the argument of perigee of theorbits of each of the plurality of satellites is in the range of 195degrees to 345 degrees for apogees in the northern hemisphere and in therange of 15 degrees to 165 degrees for apogees in the southernhemisphere.
 5. A system according to claim 1, wherein each of theplurality of satellites has throughout its orbit an orbital height lowerthan a height necessary for geostationary orbits.
 6. A system accordingto claim 1, wherein the plurality of satellites are equally spaced inmean anomaly within their respective ground tracks.
 7. A systemaccording to claim 1, wherein the orbits of the plurality of satellitesare further configured such that the portion of the orbits during whichthe communications equipment on the satellites is enabled tocommunicate, is separated from the equatorial plane of the earth by atleast a predetermined amount.
 8. A system according to claim 1, whereinthe communications equipment on the plurality of satellites is furtherconfigured to communicate at frequencies allocated to geostationarysatellites.
 9. A system according to claim 1, wherein each of theplurality of satellites has a power system configured to generate afirst amount of power when the communications equipment on the satelliteis enabled and a second amount of power more than the first amount ofpower when the communications equipment is not enabled, to store excesspower generated when the communications equipment is not enabled, and toenable the communications equipment with both the stored excess powerand the generated first amount of power.
 10. A constellation ofsatellites, comprising: a plurality of satellites in orbits around theearth having apogees and perigees, each of the satellites havingcommunications equipment thereon configured to communicate only during apredetermined portion of the satellite's orbit proximate to apogee, theorbits of the plurality of satellites being configured to form at leasttwo ground tracks on the earth displaced from each other longitudinally,each of the ground tracks repeating daily and having a number of activearcs, each active arc corresponding to the portion of the orbit of eachsatellite during which the communications equipment on the satellite isenabled to communicate, the orbits of the plurality of satellites beingfurther configured such that at all times there are at least two of thesatellites in each of the active arcs and such that at all times each ofthe satellites in any one of the active arcs is separated by at least apredetermined angle, as observed from the earth, from each othersatellite in the same active arc and from any satellite in any otheractive arc.
 11. A constellation according to claim 10, wherein the orbitof each of the plurality of satellites has a mean motion that is one of2, 3 and
 4. 12. A constellation according to claim 10, wherein the orbitof each of the plurality of satellites is inclined at criticalinclination.
 13. A constellation according to claim 10, wherein theargument of perigee of the orbits of each of the plurality of satellitesis in the range of 195 degrees to 345 degrees for apogees in thenorthern hemisphere and in the range of 15 degrees to 165 degrees forapogees in the southern hemisphere.
 14. A constellation according toclaim 10, wherein each of the plurality of satellites has throughout itsorbit a orbital height lower than a height necessary for geostationaryorbits.
 15. A constellation according to claim 10, wherein thesatellites in each of the two or more ground tracks are equally spacedin mean anomaly.
 16. A constellation according to claim 10, wherein theorbit of each of the plurality of satellites is further configured suchthat the portion of the orbits during which the communications equipmenton the satellites is enabled to communicate, is separated from theequatorial plane of the earth by a least a predetermined amount.
 17. Aconstellation according to claim 10, wherein the communicationsequipment on each of the plurality of satellites is further configuredto communicate at frequencies allocated to geostationary satellites. 18.A constellation according to claim 10, wherein each of the plurality ofsatellites has a power system configured to generate a first amount ofpower when the communications equipment on the satellite is enabled anda second amount of power more than the first amount of power when thecommunications equipment is not enabled, to store excess power generatedwhen the communications equipment is not enabled, and to enable thecommunications equipment with both the stored excess power and thegenerated first amount of power.
 19. A method for satellitecommunications, comprising: orbiting a plurality of communicationssatellites about the earth, the orbits having apogees and perigees; andenabling each of the plurality of communications satellites tocommunicate only during a predetermined portion of the orbits proximateto apogee; wherein the orbits of the plurality satellites form at leasttwo ground tracks on the earth displaced from each other longitudinally,each of the ground tracks repeating daily and having a number of activearcs, each active arc corresponding to the portion of the orbit of eachsatellite during which the communications equipment on the satellite isenabled to communicate; and wherein the satellites are orbited such thatat all times at least two of the satellites are in each of the activearcs and such that at all times each of the satellites in any one of theactive arcs is separated by at least a predetermined angle, as observedfrom the earth, from each other satellite in the same active arc andfrom any satellite in any other active arc.
 20. A method according toclaim 19, further comprising: configuring the orbits of each of theplurality of satellites to have a mean motion that is one of 2, 3 and 4.